With the continuously increasing demand for international and long distance air travel, there is a perceived need for a supersonic civilian transport aircraft. However, it is expected that such an aircraft would be expensive to produce so that airlines, and their customers, would be reluctant to use the aircraft unless the high cost can be offset by other factors, besides the convenience of high speed. Some of these factors include minimizing the mass of the aircraft so that it would consume less fuel, reduce the airline's cost per passenger mile, and increase the aircraft's range and payload. Moreover, the aircraft should have an extended life, thereby allowing the airlines to recoup their investment over a longer period of time.
The need for increased fuel efficiency, long range, high payload and longer life of a supersonic civilian transport aircraft imposes exacting new demands on materials of construction. For example, the fuselage, wings, and other components of the aircraft's outer shell, should be light, but also have high strength-to-weight ratio properties, calling for low density, high strength materials. Moreover, the materials should have high modulus, fatigue resistance for long life, and high thermo-mechanical endurance to withstand stresses under the high temperatures encountered during supersonic flight. From a safety standpoint, the materials should be damage-resistant and damage-tolerant, and from a preventive maintenance standpoint, the materials should provide visible signs of damage, long before actual failure.
Fabricating aircraft fuselages and exterior panels, such as wings and control surfaces from metals, such as titanium alloys, may not meet all the performance criteria for an advanced supersonic civilian aircraft. Titanium alloys have relatively high density compared to the target density for an advanced supersonic civilian aircraft, and are relatively expensive. Moreover, titanium panel sizes are limited, due to physical property constraints, so that a large aircraft would require many joined panels. An increased number of joints results in increased weight, an undesirable factor. Titanium alloys also have relatively low fatigue strength and relatively high crack-growth rates so that the life of an aircraft may not be extended to meet the criterion set for continuous service in supersonic civilian transport. Consequently, titanium alloys may not be the optimum material of choice.
As an alternative, aircraft fuselages and exterior panels could be fabricated from polymeric composites. Such composites include a thermosetting or thermoplastic polymer ("resin") matrix within which is embedded reinforcing fibers, such as carbon fibers. However, performance of these polymeric composites may change with time upon repeated exposure to the high temperatures encountered during supersonic flight. Such temperatures clearly vary depending upon the speed of flight, for example temperatures of up to about 350.degree. F. (about 175.degree. C.) are expected at mach 2.4. Polymeric composites are also susceptible to undetectable mechanical damage which may compromise structural integrity, and which requires additional material, to compensate for unknown risks thereby increasing aircraft mass. Furthermore, polymeric composites are also susceptible to damage from lightning strikes and therefore require additional conductive structure for protection. This also adds mass to the aircraft.
Prior art attempts at developing hybrid laminates that include layers of polymeric composite and layers of metal have not produced composites with the requisite combination of low density and physical properties necessary for use in a fuselage or exterior skin panels of a supersonic civilian aircraft. A laminate must meet the strength, modulus, fatigue resistance, and thermo-mechanical endurance properties discussed above. It should also have enhanced damage tolerance, and should desirably dent in a manner similar to metals thereby allowing detection of damage before significant physical property deterioration occurs. The polymeric composite layers in the laminate should be protected from thermal-induced oxidation, water ingress, and potential damage that could be caused by exposure to fuel and other solvents. Moreover, the laminate should exhibit high strength, and resist propagation of cracks, even from those points where it has been drilled-through to receive fasteners. The hybrid laminate should also be compatible with fusing to a core structure that may form part of the laminate structure, such as an aircraft panel.